With reference to FIG. 1, a ducted fan gas turbine engine generally indicated at 10 has a principal and rotational axis X-X. The engine comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, and intermediate pressure turbine 17, a low-pressure turbine 18 and a core engine exhaust nozzle 19. A nacelle 21 generally surrounds the engine 10 and defines the intake 11, a bypass duct 22 and a bypass exhaust nozzle 23.
The gas turbine engine 10 works in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the intermediate pressure compressor 14 and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 13 compresses the air flow A directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
The compressed air exhausted from the high-pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines 16, 17, 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines respectively drive the high and intermediate pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.
The compressor rotor blades are surrounded by a compressor casing. A small gap, or clearance, is provided radially between the tips of the compressor rotor blades and the compressor rotor casing. The compressor casing is provided with an abradable liner on its radially inner surface immediately around the tips of the compressor rotor blades. These abradable liners wear preferentially relative to the material of the tips of the compressor rotor blades during engine service. The abradable liners reduce over tip leakage between the tips of the compressor rotor blades and the compressor casing and hence reduce the associated loss in engine efficiency and engine performance.
Abradable liners are also used on the turbine casing, which is arranged radially around stages of turbine rotor blades. Abradable liners are also used on other components of gas turbine engines where there is a requirement to form a seal between a rotatable or movable component and a static component.
Conventional abradable liners are typically thermally sprayed (eg plasma sprayed) metal coatings for compressors, and thermally sprayed ceramic coatings for turbines. These coatings have to provide a balance in mechanical properties between a requirement to be soft enough to be abraded and hard enough to resist erosion. Thus thermally sprayed abradables generally have a metallic matrix (eg based on Ni, Cr or Al) and a dislocator phase to impart improved cutting characteristics.
However, there are disadvantages associated with thermally sprayed abradables. in particular, the coatings are generally limited to thicknesses of about 2 to 3 mm, as increased thicknesses can increase susceptibility to cracking and spalling due to thermal expansion coefficient mismatches and residual stresses. Other disadvantages include: difficulties with process control which can lead to coating variability and defects, resulting in blade damage or erosive abradable failures; and significant powder feedstock wastage (typically about 50%) during spraying.
An alternative to thermally sprayed abradables consists of a liner formed from a mesh of sintered fine iron-based wires brazed directly to a casing. Such liners are commonly used at operational temperatures of greater than 350° C. However, the precise structure of the mesh can be difficult to control during manufacture of the liner. Further, in use, heat can be generated through friction between the rotating member and the mesh.